As a result of increasing fuel prices during the 1970's, aircraft engine designers have sought to improve the efficiency of their product. One area of the gas turbine engine which has been studied is the compressor. Basically, the compressor consists of a number of bladed compressor disks which rotate at high speed and increase the pressure of an air stream flowing through the compressor. The high pressure air exiting the compressor is mixed with fuel and burned in a combustor. The exhaust gases are then expanded through a turbine wheel where work is extracted from the flow stream.
The airflow through the compressor can be divided into two broad regions--the endwall flow region near both the casing and the hub where viscous boundary layer effects and blade/vane tip effects dominate and the center-flow region in the central portion of the compressor where the aforementioned effects are small or negligible. Roughly 50% of all compressor loss occurs in the endwall region.
One condition which contributes to this loss, thereby reducing compressor efficiency, is caused by the gap that normally is between the end of a compressor blade and the surrounding casing in the endwall region. Air which is compressed by the rotating blade has a tendancy to backflow, or leak, over the rotor tip through this gap resulting in a tip clearance vortex. This vortex interacts with the casing wall boundary layer and produces tip loss.
The typical approach for controlling this leakage has been to minimize the clearance between the rotor tip and the surrounding casing. However, both the compressor casing and the compressor blade grow radially during periods of engine operation. In order to avoid contact between the blades and the casing, sufficient clearance must be left during normal engine operation to allow for differential growth during transient operating conditions. An alternative approach is to anticipate rubs by providing either an abradable strip in the casing or an abradable tip on the rotor blade to permit some degree of a controlled rub.
Another technique for reducing leakage across blade tips has been found to form a recess in the wall of the casing and to extend the rotor blade to be nearly line-on-line with the original casing wall. Suc recesses may accept the rotor blade tip during some or all periods of engine operation. The transition region from compressor casing to recess is typically characterized by an abrupt change from the smooth casing wall. These abrupt transition regions occur both in the forward and aft ends of the recess. For example, trenches with rectangular cross section are known wherein the transition regions are formed by right angles. Test results indicate that such trenches may provide, at best, a marginal improvement in efficiency and, under certain conditions, actually degrade performance.